Staged combustion liquid rocket engine cycle with the turbopump unit and preburner integrated into the structure of the combustion chamber

ABSTRACT

Devices and methods of rocket propulsion are disclosed. In one aspect, a staged combustion liquid rocket engine with preburner and turbopump unit (TPU) integrated into the structure of the combustion chamber is described. An initial propellant mixture is combusted in a preburner combustion chamber formed as an annulus around a main combustion chamber, the combustion products from the preburner driving the turbine of the TPU and subsequently injected into the main combustion chamber for secondary combustion along with additional propellants, generating thrust through a supersonic nozzle. The preburner inner cylindrical wall is shared with the outer cylindrical wall of the engine&#39;s main combustion chamber and the turbine is axially aligned with the main combustion chamber. Liquid propellants supplied to the engine are utilized for regenerative cooling of the combustion chamber and preburner, where the liquid propellants are gasified in cooling manifolds before injection into the preburner and main combustion chamber.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/013,457 filed Sep. 4, 2020 and titled “Staged Combustion LiquidRocket Engine Cycle with the Turbopump Unit and Preburner Integratedinto the Structure of the Combustion Chamber,” which in turn claims thebenefit of priority to U.S. Provisional Patent Application Nos.62/897,044, filed Sep. 6, 2019 and titled “Method of Operation ofMethane LRE with the TPU Drive Gas Generator Integrated into theStructure of the Combustion Chamber,” and 62/724,580, filed Aug. 29,2018 and titled “A Method of Operation of Methane LRE with the TPU DriveGas Generator Integrated into the Structure of the Combustion Chamber,”the disclosures of which are hereby incorporated herein by reference inentirety.

FIELD

The disclosure relates generally to devices and methods of rocketpropulsion, and more specifically to a staged combustion bipropellantLiquid Rocket Engine (LRE) with the Turbo Pump Unit (TPU) and preburnerintegrated into the structure of the main combustion chamber.

BACKGROUND

Liquid Rocket Engines (LRE) are primarily used in the aerospace industryto power launch vehicles into space. Development of new LREs seek toimprove efficiency in any of several ways. The disclosure improves theoverall LRE efficiency through combining components in a manner thatreduces the overall weight of the LRE and simplifies the method ofoperation of the high efficiency staged combustion cycle.

In one embodiment, a unique and non-obvious combination and interactionbetween an LRE main combustion chamber, LRE preburner, and LRE TurboPump Unit (TPU) is disclosed. Generally, the disclosure addresses, amongother things, the task of ensuring the operability, compactness, drymass reduction and efficiency of the LRE by combusting an initialpropellant mixture in a preburner chamber which forms an annulus aroundthe main combustion chamber of the engine as a single structural unit,where the preburner inner cylindrical wall is shared with the outercylindrical wall of the engine's main combustion chamber. Aftercombustion in the annular preburner chamber, the combusted initialpropellant mixture drives the turbine of the TPU and is subsequentlyinjected into the main combustion chamber for secondary combustion alongwith additional propellants. At the same time, the LRE is configuredwith an elegantly simple design based on modern science andmanufacturing technology development; the result is a highly integratedengine that requires relatively minimal funds and time to produce as aserial model using additive manufacturing techniques. Furthermore, theengine is reliable, restartable, and provides optimal masscharacteristics while implementing preset technical parameters inspecific design embodiments.

SUMMARY

The present disclosure can provide a number of advantages depending onthe particular aspect, embodiment, and/or configuration.

Devices and methods of rocket propulsion are disclosed. In one aspect, amethane Liquid Rocket Engine (LRE) with the TPU drive gas generatorintegrated into the structure of the combustion chamber is described.

In one embodiment, a liquid rocket engine (LRE) is disclosed, the LREcomprising: a fuel supply line containing a fuel; an oxidizer supplyline containing an oxidizer; an oxidizer throttle receiving the oxidizerfrom the oxidizer supply line and splitting the oxidizer supply lineinto a first partial oxidizer supply line, a second partial oxidizersupply line, and a third partial oxidizer supply line; a turbine mountedto a turbine shaft along a longitudinal centerline of the LRE at alongitudinal proximal location of the LRE, the turbine rotating with theturbine shaft about the longitudinal centerline; a gas duct positionedlongitudinally distal to the turbine and in fluid communication with theturbine; a main combustion chamber positioned longitudinally distal tothe gas duct, having an injector head in fluid communication with boththe gas duct and the second partial oxidizer supply line; a nozzlepositioned longitudinally distal to the main combustion chamber and influid communication with the main combustion chamber; a preburnercombustion chamber positioned axially distal to the main combustionchamber, having a preburner injector head in fluid communication withboth the first cooling manifold and second cooling manifold; a firstcooling manifold positioned axially distal to the nozzle and in fluidcommunication with the fuel supply line; a second cooling manifoldpositioned axially between the preburner combustion chamber and the maincombustion chamber, the second cooling manifold in fluid communicationwith the first partial oxidizer supply line; and a third coolingmanifold positioned longitudinally distal to the preburner combustionchamber and in fluid communication with the third partial oxidizersupply line and a nozzle throat of the nozzle; wherein: the fuelreceived by the first cooling manifold from the fuel supply line becomesa gasified fuel during transport due to thermal energy transfer from thenozzle, the gasified fuel is supplied to the preburner injector head;the oxidizer received by the second cooling manifold from the firstpartial oxidizer supply line becomes a gasified oxidizer due to thermalenergy received from at least one of the gas duct, the main combustionchamber, and the preburner combustion chamber, the gasified oxidizersupplied to the preburner injector head; the preburner injector headinjects the gasified fuel with the gasified oxidizer to create preburnercombustion exhaust products which flow from the preburner combustionchamber to the turbine and drive the turbine about the turbine shaft,the preburner combustion exhaust products flowing from the turbine tothe gas duct; and the main combustion chamber injector head injects thepreburner combustion exhaust products received from the gas duct withoxidizer received from the second partial oxidizer supply line toproduce an LRE thrust directed through the nozzle exit.

In one aspect, the fuel is methane. In another aspect, the maincombustion chamber comprises a main chamber ignitor head which ignitesthe preburner combustion exhaust products with the oxidizer. In anotheraspect, the turbine is a centrifugal turbine. In another aspect, thepreburner combustion chamber comprises a set of flow vanes, the set offlow vanes directing the preburner combustion exhaust products to a setof blades of the turbine. In another aspect, the gasified fuel suppliedto the preburner injector head is completely gasified. In anotheraspect, the LRE forms a closed propellant supply scheme LRE. In anotheraspect, the preburner combustion chamber and the second cooling manifoldat least partially share a common surface side wall. In another aspect,the third manifold is positioned at a throat of the nozzle. In anotheraspect, the fuel of the fuel supply line is provided by a fuel pump, thefuel pump driven by the turbine shaft. In another aspect, the oxidizerof the oxidizer supply line is provided by an oxidizer pump, theoxidizer pump driven by the turbine shaft. In another aspect, thepreburner combustion chamber forms an annulus around the main combustionchamber. In another aspect, cooling manifold three is configured toprovide at least one of a film and a transpiration cooling of the throatof the nozzle.

In another embodiment, a method of operating a liquid rocket engine(LRE) is disclosed, the method comprising: providing a LRE comprising: afuel supply line containing a fuel; an oxidizer supply line containingan oxidizer; an oxidizer bypass regulator receiving the oxidizer fromthe oxidizer supply line and splitting the oxidizer supply line into afirst partial oxidizer supply line, a second partial oxidizer supplyline, and a third partial oxidizer supply line; a turbine mounted to aturbine shaft along a longitudinal centerline of the LRE at an LRElongitudinal proximal location, the turbine rotating with the turbineshaft; a gas duct positioned longitudinally distal to the turbine and influid communication with the turbine; a main combustion chamberpositioned longitudinally distal to the gas duct and in fluidcommunication with the gas duct, the main combustion chamber having aninjector head in fluid communication with the second partial oxidizersupply line; a nozzle positioned longitudinally distal to the maincombustion chamber and in fluid communication with the main combustionchamber; a preburner combustion chamber positioned axially distal to themain combustion chamber; a first cooling manifold positioned axiallydistal to the nozzle and in fluid communication with the fuel supplyline; a second cooling manifold positioned axially between the preburnercombustion chamber and the main combustion chamber, the second coolingmanifold in fluid communication with the first partial oxidizer supplyline and the preburner combustion chamber; and a third cooling manifoldpositioned longitudinally distal to the preburner combustion chamber andin fluid communication with the third partial oxidizer supply line andthe preburner combustion chamber; gasifying the fuel received in thefirst cooling manifold by way of heat transfer from the nozzle to thefirst cooling manifold to produce gasified fuel; gasifying the oxidizerreceived by the second cooling manifold by way of heat transfer from thegas chamber and the main combustion chamber to the second coolingmanifold to produce gasified oxidizer; igniting the gasified fuel withthe gasified oxidizer at the preburner ignitor head to produce preburnercombustion chamber exhaust products; passing the preburner combustionchamber products to the main combustion chamber by way of the gas duct;and igniting the preburner combustion chamber products with the oxidizersupplied by the second partial oxidizer supply fuel line to produce anLRE thrust directed through the nozzle.

In one aspect, the fuel is methane. In another aspect, the fuel of thefuel supply line is provided by a fuel pump, the fuel pump driven by theturbine shaft; and the oxidizer of the oxidizer supply line is providedby an oxidizer pump, the oxidizer pump driven by the turbine shaft. Inanother aspect, the preburner combustion chamber forms an annulus aroundthe main combustion chamber. In another aspect, the turbine is acentrifugal turbine. In another aspect, the preburner combustion chamberand the second cooling manifold at least partially share a commonsurface side wall. In another aspect, the gasified fuel supplied to thepreburner injector head is completely gasified; the cooling manifoldthree is configured to provide at least one of a film and atranspiration cooling of a throat of the nozzle; and the LRE forms aclosed propellant supply scheme LRE.

In another aspect, the preburner injector head injects the gasified fuelwith the gasified oxidizer into the preburner combustion chamber so asto thoroughly mix the fuel and oxidizer together and to disperse them inthe combustion chamber where they are ignited and combusted to createpreburner combustion exhaust products which flow from the preburnercombustion chamber to the turbine and drive the turbine about theturbine shaft, the preburner combustion exhaust products flowing fromthe turbine to the gas duct.

In yet another aspect, the main combustion chamber injector head injectsthe preburner combustion exhaust products received from the gas ductwith oxidizer received from the second partial oxidizer supply line into the main combustion chamber so as to thoroughly mix the fuel andoxidizer together and to disperse them in the combustion chamber wherethey are ignited and combusted to produce a fluid flow which is chokedand expanded at supersonic velocity through the nozzle, generatingthrust directed through the nozzle exit.

The phrases “at least one”, “one or more”, and “and/or” are open-endedexpressions that are both conjunctive and disjunctive in operation. Forexample, each of the expressions “at least one of A, B and C”, “at leastone of A, B, or C”, “one or more of A, B, and C”, “one or more of A, B,or C” and “A, B, and/or C” means A alone, B alone, C alone, A and Btogether, A and C together, B and C together, or A, B and C together.

The term “a” or “an” entity refers to one or more of that entity. Assuch, the terms “a” (or “an”), “one or more” and “at least one” can beused interchangeably herein. It is also to be noted that the terms“comprising”, “including”, and “having” can be used interchangeably. Theterm “automatic” and variations thereof, as used herein, refers to anyprocess or operation done without material human input when the processor operation is performed. However, a process or operation can beautomatic, even though performance of the process or operation usesmaterial or immaterial human input, if the input is received beforeperformance of the process or operation. Human input is deemed to bematerial if such input influences how the process or operation will beperformed. Human input that consents to the performance of the processor operation is not deemed to be “material”.

The terms “determine”, “calculate” and “compute,” and variationsthereof, as used herein, are used interchangeably and include any typeof methodology, process, mathematical operation or technique.

The term “means” as used herein shall be given its broadest possibleinterpretation in accordance with 35 U.S.C., Section 112, Paragraph 6.Accordingly, a claim incorporating the term “means” shall cover allstructures, materials, or acts set forth herein, and all of theequivalents thereof. Further, the structures, materials or acts and theequivalents thereof shall include all those described in the summary,brief description of the drawings, detailed description, abstract, andclaims themselves.

The term “module” as used herein refers to any known or later developedhardware, software, firmware, artificial intelligence, fuzzy logic, orcombination of hardware and software that can perform the functionalityassociated with that element.

The phrase “graphical user interface” or “GUI” means a computer-baseddisplay that allows interaction with a user with aid of images orgraphics.

The term “computer-readable medium” as used herein refers to any storageand/or transmission medium that participate in providing instructions toa processor for execution. Such a computer-readable medium is commonlytangible, non-transitory, and non-transient and can take many forms,including but not limited to, non-volatile media, volatile media, andtransmission media and includes without limitation random access memory(“RAM”), read only memory (“ROM”), and the like. Non-volatile mediaincludes, for example, NVRAM, or magnetic or optical disks. Volatilemedia includes dynamic memory, such as main memory. Common forms ofcomputer-readable media include, for example, a floppy disk (includingwithout limitation a Bernoulli cartridge, ZIP drive, and JAZ drive), aflexible disk, hard disk, magnetic tape or cassettes, or any othermagnetic medium, magneto-optical medium, a digital video disk (such asCD-ROM), any other optical medium, punch cards, paper tape, any otherphysical medium with patterns of holes, a RAM, a PROM, and EPROM, aFLASH-EPROM, a solid state medium like a memory card, any other memorychip or cartridge, a carrier wave as described hereinafter, or any othermedium from which a computer can read. A digital file attachment toe-mail or other self-contained information archive or set of archives isconsidered a distribution medium equivalent to a tangible storagemedium. When the computer-readable media is configured as a database, itis to be understood that the database may be any type of database, suchas relational, hierarchical, object-oriented, and/or the like.Accordingly, the disclosure is considered to include a tangible storagemedium or distribution medium and prior art-recognized equivalents andsuccessor media, in which the software implementations of the presentdisclosure are stored. Computer-readable storage medium commonlyexcludes transient storage media, particularly electrical, magnetic,electromagnetic, optical, magneto-optical signals.

Moreover, the disclosed methods may be readily implemented in softwareand/or firmware that can be stored on a storage medium to improve theperformance of: a programmed general-purpose computer with thecooperation of a controller and memory, a special purpose computer, amicroprocessor, or the like. In these instances, the systems and methodscan be implemented as program embedded on personal computer such as anapplet, JAVA.RTM. or CGI script, as a resource residing on a server orcomputer workstation, as a routine embedded in a dedicated communicationsystem or system component, or the like. The system can also beimplemented by physically incorporating the system and/or method into asoftware and/or hardware system, such as the hardware and softwaresystems of a communications transceiver.

Various embodiments may also or alternatively be implemented fully orpartially in software and/or firmware. This software and/or firmware maytake the form of instructions contained in or on a non-transitorycomputer-readable storage medium. Those instructions may then be readand executed by one or more processors to enable performance of theoperations described herein. The instructions may be in any suitableform, such as but not limited to source code, compiled code, interpretedcode, executable code, static code, dynamic code, and the like. Such acomputer-readable medium may include any tangible non-transitory mediumfor storing information in a form readable by one or more computers,such as but not limited to read only memory (ROM); random access memory(RAM); magnetic disk storage media; optical storage media; a flashmemory, etc.

The preceding is a simplified summary of the disclosure to provide anunderstanding of some aspects of the disclosure. This summary is neitheran extensive nor exhaustive overview of the disclosure and its variousaspects, embodiments, and/or configurations. It is intended neither toidentify key or critical elements of the disclosure nor to delineate thescope of the disclosure but to present selected concepts of thedisclosure in a simplified form as an introduction to the more detaileddescription presented below. As will be appreciated, other aspects,embodiments, and/or configurations of the disclosure are possibleutilizing, alone or in combination, one or more of the features setforth above or described in detail below. Also, while the disclosure ispresented in terms of exemplary embodiments, it should be appreciatedthat individual aspects of the disclosure can be separately claimed.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure will be readily understood by the following detaileddescription in conjunction with the accompanying drawings, wherein likereference numerals designate like elements. The elements of the drawingsare not necessarily to scale relative to each other. Identical referencenumerals have been used, where possible, to designate identical featuresthat are common to the figures.

FIG. 1 shows a schematic diagram of one embodiment of a methane/oxygenLRE system with the preburner and TPU integrated into the structure ofthe combustion chamber; and

FIG. 2 shows a flowchart of one method of operation of themethane/oxygen LRE system of FIG. 1 .

It should be understood that the proportions and dimensions (eitherrelative or absolute) of the various features and elements (andcollections and groupings thereof) and the boundaries, separations, andpositional relationships presented there between, are provided in theaccompanying figures merely to facilitate an understanding of thevarious embodiments described herein and, accordingly, may notnecessarily be presented or illustrated to scale, and are not intendedto indicate any preference or requirement for an illustrated embodimentto the exclusion of embodiments described with reference thereto.

DETAILED DESCRIPTION

Reference will now be made in detail to representative embodiments. Thefollowing descriptions are not intended to limit the embodiments to onepreferred embodiment. To the contrary, it is intended to coveralternatives, modifications, and equivalents as can be included withinthe spirit and scope of the described embodiments as defined, forexample, by the appended claims.

FIG. 1 is a schematic diagram of one embodiment of a methane/oxygen LREsystem with the preburner and TPU integrated into the structure of thecombustion chamber (also referred to as “methane/oxygen LRE system,”“LRE system,” and/or simply “LRE”). FIG. 2 is a flowchart of one methodof operation of the methane/oxygen LRE system (also referred to as “LREsystem method” or “LRE method”) of FIG. 1 .

Generally, an LRE is disclosed in which combustion of an initialpropellant mixture is performed in a preburner chamber, the preburnerchamber forming an annulus around a main combustion chamber of theengine as a single structural unit. The preburner inner cylindrical wallis shared with the outer cylindrical wall of the engine's maincombustion chamber and contains a part of cooling manifold two,providing cooling to both the preburner combustion chamber and maincombustion chamber. After combustion in the annular preburner chamber,the combusted initial propellant mixture drives the turbine of a turbopump unit (TPU) and is subsequently injected into the main combustionchamber for secondary combustion along with additional propellants.

In other aspects of the disclosed LRE, the amount of liquid fuelsupplied to the engine is fully utilized for regenerative cooling of thenozzle part of combustion chamber, where the liquid fuel is completelygasified and enters the cavities of a preburner where it is partiallyburnt in the preburner combustion chamber. The liquid oxidizer issupplied partly to the combustion chamber directly, and partly, withpartial gasification, to the fuel-rich gas generator, where the gasifiedfuel that was supplied is burnt together with the oxidizer. Theresulting fuel-rich gas is principally used to power a centrifugalturbine; afterwards the fuel-rich gas is burnt again in the combustionchamber of the engine with additional oxidizer generating thrust througha supersonic nozzle. This described LRE is a closed propellant supplyscheme engine.

With attention to FIG. 1 , the Liquid Rocket Engine (LRE) 100 comprisesan upper or proximal LRE portion 101 and a lower or distal portion 102.A longitudinal axis 110 passes through and extends pass the proximal LREportion 101 and the distal portion 102. For example, the longitudinalaxis 110 extends from distal LRE portion 102 through nozzle 110, asdepicted in FIG. 1 .

A series or sequence of components of the LRE 100 are assembled in alinear fashion along the longitudinal axis 110. Moving from theuppermost or most proximal location downwards or distally along thelongitudinal axis 110, a sequence of the following components arepositioned: fuel pump 2, oxidizer pump 6, turbine 20, gas duct 21, maincombustion chamber 23, and nozzle 24. Also, positioned axiallysurrounding the gas duct 21 and at least substantially most of both ofthe turbine 20 and main combustion chamber 23, are each of the preburnercombustion chamber 18 and the cooling manifold two 13. A TPU shaft 1runs along the longitudinal axis 110, the TPU shaft 1 positioned at anaxial centerline of the fuel pump 2, oxidizer pump 6, turbine 20, gasduct 21, main combustion chamber 23, nozzle 24, preburner combustionchamber 18 and the cooling manifold two 13.

Fuel pump 2 receives fuel from a fuel tank (not shown) by way of fuelfeed line 210 and is driven by turbine 20 by way of TPU shaft 1. Thefuel pump 2 supplies fuel to fuel supply line 3 (note: ΔP_(FL)). Thefuel pump may be referred to as FP. The fuel pump 2 provides fuel to theLRE. The fuel supply line 3 runs to, or is in fluid communication with,a fuel bypass regulator valve 4 (note FBR and ΔP_(r.f)), which canreroute fuel upstream of the fuel pump 2. The fuel supply line 3 alsoruns to, or is in fluid communication with, a line regulator value (noteFLR). The fuel supply line 3 runs to, or connects with, cooling manifoldone 11 so as to supply fuel to the cooling manifold one 11. As fuelprovided by fuel supply line 3 flows through cooling manifold one 11 thefuel becomes gasified wherein the gasified fuel is received at preburnerinjector head 17.

The term “fluid” means a substance devoid of shape and yields toexternal pressure, to include liquids and gases, e.g. fuels or oxidizersin liquid or gaseous form). The phrase “fluid communication” means afluid flows or runs between or within two or more locations, elements,or components.

Oxidizer pump 6 receives oxidizer from an oxidizer tank (not shown) byway of oxidizer feed line 610 and is driven by turbine 20 by way of TPUshaft 1. The oxidizer pump 6 supplies oxidizer to main oxidizer supplyline 7 (note ΔP_(OL)). The oxidizer pump 6 may be referred to as OP. Theoxidizer pump 6 provides oxidizer to the LRE at three positions. Anoxidizer bypass regulator valve 8 (note OBR and ΔP_(r.o)) reroutes fuelupstream of the oxidizer pump 6. (Note that “downstream” refers to aposition relatively forward in a fluid flow, meaning a position in thedirection of the fluid flow, and “upstream” refers to a positionrelatively behind in a fluid flow, meaning a position in the oppositedirection of the fluid flow.) After, or downstream, of the oxidizerregular value 8, the oxidizer supply line 7 continues to the oxidizerline regulator valve 9 (note OLR), then continues to the oxidizerthrottle 10.

In one embodiment, the oxidizer pump 6 and/or fuel pump 2 form componentloop system for engine control.

The oxidizer throttle 10 splits the oxidizer supply line 7 into threepartial oxidizer supply lines: a first partial oxidizer supply line 12,a second partial oxidizer supply line 16, and a third partial oxidizersupply line 14.

The first partial oxidizer supply line 12 supplies oxidizer to coolingmanifold two 13 (note ΔP_(OL) _(gg) ) with gasification coefficientβ_(o)<1.0. Note that β is the fractional amount of a fluid which is in agaseous state (1−β is the amount in a liquid state). The subscripts oand f indicate oxidizer (usually oxygen) and fuel (usually methane oranother hydrocarbon).

With respect to the cooling manifold two 13, a gasification coefficientβ_(o)<1.0 means that the oxidizer, as supplied on entry to the coolingmanifold two 13 from first partial oxidizer supply line 12, is not fullygasified (as fully gasified would have a gasification coefficientβ_(o)=1.0). Note that as the oxidizer travels or flows within coolingmanifold two 13, the oxidizer receives thermal energy (e.g. heat) fromthe adjacent LRE components (e.g. from one or more of turbine 20, gasduct 21, main combustion chamber 23) which heats the oxidizer andincreases the oxidizer's gasification, meaning β_(o) increases in value.Stated another way, as the oxidizer provided by first partial oxidizersupply line 12 flows through cooling manifold two 13, the oxidizerbecomes gasified (wherein the gasified fuel is received at preburnerinjector head 17). Also, as the oxidizer travels or flows within coolingmanifold two 13, the receipt of heat from adjacent LRE components servesto cool such components.

The second partial oxidizer supply line 16 supplies oxidizer to injectorhead 22 of the main combustion chamber 23. The oxidizer supplied bysecond partial oxidizer supply line 16 to injector head 22 hasgasification 1−β_(o).

The third partial oxidizer supply line 14 supplies oxidizer to coolingmanifold three 15. The cooling manifold three 15 operates with ΔP_(OL)_(col) .

Generally, various fluid supply lines (e.g. first partial oxidizersupply line 12 supplying oxidizer to cooling manifold two 13 and fuelline 3 supplying fuel to cooling manifold 11) are positioned so as toreceive heat from LRE components, thereby providing a cooling to thosecomponents by receiving heat from those components. The received heatalso serves to gasify the fluid operating within a particular oxidizersupply line. The terms “gasify,” or “gasifying,” and “gasification” meanto increase the percentage of gas relative to liquid in a fluid, i.e. toincrease the value of (1−β_(o)).

Gasified fuel or gasified oxidizer is produced by using respectiveliquid fuel or oxidizer for cooling. When a fuel or an oxidizer is fullygasified, its effective cooling ability has been consumed.

Cooling manifold one 11 (note ΔP_(FL) _(col) ) is positioned adjacentnozzle 24 and receives fuel from fuel supply line 3. The supply line 3provides or supplies fuel at a lower or distal position on the nozzle24, meaning at a distal position of the LRE and at a distal position ofthe longitudinal axis 110. The received fuel from supply line 3 thentravels upwards (or generally from a distal LRE location to a proximalLRE location) along or adjacent the nozzle 24, receiving thermal energysuch as heat from the nozzle 24. The received heat serves to cool thenozzle 24 and also to gasify the fuel. Upon reaching the preburnerinjection head 17, after traveling the length of cooling manifold one11, the fuel is, in one embodiment, fully gasified and thus has agasification coefficient β_(f)=1.0 at exit. In one embodiment, thegasified fuel is then supplied to the preburner injector head 17. Uponreaching the upon reaching the preburner injection head 17, the fuel issubstantially fully gasified, meaning the fuel has a gasificationcoefficient β_(f)˜1.0 at exit.

The preburner injection head 17 mixes and blends gasified fuel receivedfrom the cooling manifold one 11 and gasified oxidizer received fromcooling manifold two 13. Note that a set of preburner injection heads 17are positioned at a lower or bottom or distal position of the coolingmanifold two 13 and preburner combustion chamber 18 and may be axiallysymmetrically positioned about the LRE longitudinal centerline 110.

Adjacent to the mixing or combination of fluids at the preburnerinjection head 17 is preburner injector flow 171, where fluid flow movesbetween the cooling manifold two 13 and preburner combustion chamber 18.

The TPU shaft 1 is formed along and rotates about the LRE longitudinalcenterline 110. Note N_(FP) and N_(OP) associated with the TPU shaft.The TPU shaft 1, driven by turbine 20, operates each of the fuel pump 2and oxidizer pump 6 through supply of rotational energy or rotationalpower. The TPU shaft 1 is positioned in the proximal LRE portion 101area.

The turbine 20 is positioned below or distal to the TPU shaft 1 and isdriven by preburner exhaust products received from the preburnercombustion chamber 18. The preburner exhaust products are guided ordirected to the turbine 20 by way of a set of flow vanes 19. Note CT andN_(T) associated with turbine 20. In one embodiment, the turbine 20 is acentrifugal turbine.

In other embodiments, the turbine 20 is any turbine known to thoseskilled in the art that may function in an LRE environment. Fluidleaving the turbine moves or travels downstream and is received by thegas duct 21. The turbine 20 is positioned and configured to be axiallysymmetric about the LRE longitudinal centerline 110.

Gas duct 21 (note ΔP_(GD)) is fitted to receive fluid from the turbine20 and pass the fluid through the main combustion chamber injector head22 to the main combustion chamber 23. The fluid operating within gasduct is the preburner combustion (exhaust) products passed from orreceived from the preburner combustion chamber 18 by the turbine 20 andin turn passed to and received by the gas duct 21. The gas duct 21 maybe configured with one or more flow straighteners (not shown) to reduceturbulence from the turbine 20. The gas duct 21 may generally beconfigured as a converging/diverging nozzle. The gas duct 21 ispositioned and configured to be axially symmetric about the LRElongitudinal centerline 110. The gas duct 21 provides fluid to the maincombustion chamber. The fluid passed or provided from the gas duct 21 tothe main combustion chamber 22 is the preburner combustion (exhaust)products that was received by the turbine 20 from the preburnercombustion chamber 18.

The main combustion chamber injector 22 receives fluid from the gaschamber 21. The main combustion chamber 23 may have a proximal portionof narrower diameter than the body of the main combustion chamber 23 toenable, among other things, fitting to the gas chamber. The maincombustion chamber 23 has a main combustion chamber injector head 22positioned at an upper or proximal position of the main combustionchamber 23. The main combustion chamber injector head 22 has aΔP_(o.inj). The main chamber injector head 22 receives oxidizer fromsecond partial oxidizer supply line 16 and combines the receivedoxidizer with the preburner combustion (exhaust) products received fromthe gas duct 21. Combustion of these fluids occurs in the combustionchamber 23, resulting in an LRE generated thrust directed downwards ordistally through the nozzle 24.

The main combustion chamber 23 is positioned and configured to beaxially symmetric about the LRE longitudinal centerline 110. The maincombustion chamber injector head 22 may be positioned and configured tobe axially symmetric about the LRE longitudinal centerline 110.

The nozzle 24 is a supersonic nozzle. The nozzle 24 is configured toreceive the cooling manifold one 11 about the nozzle exterior andconfigured to receive the cooling manifold three 15 adjacent to and/ornear a neck of the nozzle 24. Thermal energy, e.g. heat energy, ispassed from the nozzle 24 to the cooling manifold one 11, wherein fuelsupplied by fuel supply line 3 is gasified. The small portion ofoxidizer contained in cooling manifold three 15 is injected into thelower portion of the combustion chamber 23 above and along the nozzlethroat via transpiration, film, or other injection cooling method. Thenozzle 24 is positioned and configured to be axially symmetric about theLRE longitudinal centerline 110.

Cooling manifold two 13 is positioned axially between the turbine 20,gas duct 21, and main combustion chamber 23. More specifically, thecooling manifold two 13 is positioned axially between the maincombustion chamber outer wall 232 and the preburner inner wall 181. Asprovided above, the cooling manifold two 13 receives oxidizer fluid fromthe first partial oxidizer supply line 12, the fluid upon receipt by thecooling manifold two 13 not fully gasified. As the received fluid flowswithin cooling manifold two 13, the oxidizer fluid receives thermalenergy (e.g. heat) from the adjacent LRE components (e.g. from one ormore of turbine 20, gas duct 21, main combustion chamber 23) which heatsthe oxidizer and increases the oxidizer's gasification: the receivedoxidizer fluid becomes gasified during flow through the cooling manifoldtwo 13. The gasified oxidizer fluid operating within the coolingmanifold two 13 flows downward or distally within the LRE toward thepreburner injector head 17 and the preburner injector flow 171. At thepreburner injector flow 171, fluid flow from the cooling manifold two 13moves to the preburner combustion chamber 18. The cooling manifold two13 may be positioned and configured to be axially symmetric about theLRE longitudinal centerline 110.

The preburner combustion chamber 18 (note ΔP_(o.gas) _(gg) andΔP_(F.gas) _(gg) ) is positioned axially distal to the cooling manifoldtwo 13 and has preburner inner wall 181 and preburner outer wall 182.The preburner combustion chamber 18 forms an annular cavity.

The preburner combustion chamber 18 combusts oxidizer and fuel ignitedby and at the preburner injector head 17, specifically gasified fuelreceived from the cooling manifold one 11 and gasified oxidizer receivedfrom cooling manifold two 13. The combustion of fluids by the preburnercombustion chamber 18 produces preburner combustion exhaust productswhich are passed or provided to the turbine 20, resulting in rotationaloperation of the turbine 20. The preburner combustion chamber 18 isassociated with ΔP_(o.gas) _(gg) and ΔP_(F.gas) _(gg) . The preburnerCombustion Chamber 18 may include a set of flow vanes 19 to align thecombusted preburner exhaust flow into the turbine 20. The preburnerCombustion Chamber 18 and/or flow vanes 19 may be positioned andconfigured to be axially symmetric about the LRE longitudinal centerline110.

Cooling manifold three 15 may be configured with transpiration or filmcooling holes, or another injection cooling mechanism, to cool thenozzle throat. The nozzle throat is typically the hottest area in an LREand may require additional cooling to prevent overheating and damage tothe nozzle. In one embodiment, the cooling manifold three 15 isconfigured to provide film or transpiration cooling of the nozzle throat24.

Both the preburner injector 17 and the main injector 22 serve toseparate propellants in many small feed lines which, during injectioninto the respective combustion chamber, induce mixing, dispersion, and(if any liquid is injected) atomization of the propellant for efficientand complete combustion. In the case of the preburner, “completecombustion” means only complete combustion of the oxidizer only withpartial combustion of the fuel.

The performance of the LRE 100 of FIG. 1 may be determined by thenontrivial solution to the power balance equation N_(T)=N_(ΣP) _(i)where N_(T) is the power generated by the turbine (20), driven by thecombustion in the preburner (18), and N_(ΣP) _(i) is the power requiredby the turbine pumps (2 and 6) to overcome the sum of hydraulic lossesthroughout the propellant supply system (3-5 and 7-9), which isdependent upon the selected pressure of the main combustion chamber(23). Here,

$N_{T} = {\eta_{t}{\overset{.}{m}}_{gg}R_{gg}T_{gg}\frac{k_{gg}}{k_{gg} - 1}\left( {1 - \left( \delta_{g}^{\frac{k_{gg} - 1}{k_{gg}}} \right)^{- 1}} \right)}$

where η_(t) is the turbine efficiency, {dot over (m)}_(gg) is mass flowrate through the preburner, k_(gg) is the adiabatic index of combustion,R_(gg) is the gas constant, T_(gg) is the combustion temperature, andδ_(g) is the pressure drop ratio in the preburner. The subscript ggstands for gas generator, referring to the preburner combustor. Also,

$N_{\sum P_{i}} = {{\frac{{\overset{.}{m}}_{f}\eta_{cf}}{\eta_{fp}\rho_{f}}\left\lbrack {{\frac{\delta_{g}}{\eta_{{gm}_{T}}}\left( {P_{ch} + {\Delta P_{gd}}} \right)} + {\Delta P_{f.l}} - P_{{in}.{fp}}} \right\rbrack} + {\frac{{\overset{.}{m}}_{o}\eta_{co}}{\eta_{op}\rho_{o}}\left\lbrack {{\frac{\delta_{g}}{\eta_{{gm}_{T}}}\left( {P_{ch} + {\Delta P_{gd}}} \right)} + {\Delta P_{o.l}} - P_{{in}.{op}}} \right\rbrack}}$where each of the two large sets of brackets refers to the sameparameters of the fuel feed system (subscript f, first brackets) and theoxidizer feed system (subscript o, second brackets). {dot over (m)}_(x)is the total mass flow of a propellant, η_(cx) is the pump losscoefficient, η_(xp) is the pump efficiency, ρ_(x) is the fluid density,η_(gm) _(T) is the turbine loss coefficient, P_(ch) is the maincombustion chamber pressure, ΔP_(gd) is the pressure drop in the gasduct, ΔP_(x.l) is the total pressure loss across a propellant feedsystem, and P_(in.xp) is the pressure at the pump inlet. (note that xmay be f if relative to fuel and may be o if relative to oxidizer.)

In the embodiment disclosed in FIG. 1 , all fuel (methane) is combustedin the preburner with a part of oxidizer (oxygen), then recombusted withadditional oxidizer in the main combustion chamber. This is calledFuel-Rich Staged Combustion.

An alternative embodiment of the LRE 100 routes all oxidizer through thepreburner 18 with a part of fuel, followed by secondary combustion withadditional fuel in the main combustion chamber 23. This alternativemethod is called an Oxidizer-Rich Staged Combustion and is more amenableto kerosene fuel. In this embodiment, nearly 100% of oxidizer is fed tocooling manifold one 11, instead of 100% of fuel, with a nearlynegligible amount of oxidizer still being fed to cooling manifold three15 for film or transpiration cooling. The fuel is instead fed in part tocooling manifold two 13 and in part directly to the main combustionchamber injector head 22 through a supply line equivalent to the mainoxidizer supply line 16 in the fuel-rich version of the cycle.

FIG. 2 presents one method of regular or steady-state operation of themethane/oxygen LRE system of FIG. 1 (as opposed to transitional states,such as start-up). In one embodiment of a method of use, the LRE system100 of FIG. 1 follows the sequence of steps described in FIG. 2 . Othermethods of use are possible, to include a sequence of steps differentthan those of FIG. 2 , a sequence with additional steps, and a sequencewith fewer steps. Also, as will be clear from the below description,elements of the LRE system 100 of FIG. 1 , and/or other aspects of anLRE system as described in this disclosure, may be incorporated.

With particular attention to FIG. 2 , a flowchart of a method of use ofan LRE in steady-state operation is provided, the method 200 utilizingthe elements described in the systems of FIG. 1 .

The method 200 starts at step 204 and ends at step 236. Any of thesteps, functions, and operations discussed herein can be performedcontinuously and automatically. In some embodiments, one or more of thesteps of the method 200 may comprise computer control, use of computerprocessors, and/or some level of automation. For example, one or morecomponents of the LRE 100, to include the entire LRE 100, may becontrolled with aid of an engine control system. The steps arenotionally followed in increasing numerical sequence, although, in someembodiments, some steps may be omitted, some steps added, and the stepsmay follow other than increasing numerical order.

At step 208, a liquid rocket engine (LRE) 100 of the type described withrespect to FIG. 1 is provided. The LRE 100 is engaged with supplementalelements not explicitly described in FIG. 1 , such as a fuel tank toprovide fuel to fuel pump 2 and an oxidizer tank to supply oxidizer tooxidizer pump 6. After the completion of step 208, the method 200continues to step 212.

At step 212, the oxidizer received at cooling manifold one 11 isgasified by way of traveling through or flowing through the coolingmanifold one 11. After the completion of step 212, the method 200continues to step 216.

At step 216, fuel received at cooling manifold two 13 is gasified by wayof traveling through or flowing through the cooling manifold two 13.After the completion of step 216, the method 200 continues to step 220.

At step 220, the gasified oxidizer produced in step 212 is combined withthe gasified fuel produced at step 216 is combusted within preburnercombustion chamber 18 with aid of the preburner injector head 17, thecombustion producing preburner combustion chamber exhaust products. Thepreburner combustion chamber exhaust products are directed, by way of aset of flow vanes 19, to the turbine 20. After the completion of step220, the method 200 continues to step 224.

At step 224, the preburner combustion chamber exhaust are used to drivethe turbine 20. The turbine drives the fuel pump 2 and the oxidizer pumpin an in-line or linear configuration or arrangement. After thecompletion of step 224, the method 200 continues to step 228.

At step 228, the preburner combustion chamber exhaust products arepassed to or flow to the gas duct 21 and then in turn to the maincombustion chamber 23. The preburner combustion chamber exhaust productsare combined or mixed with oxidizer provided by the second partialoxidizer supply line 16. The second partial oxidizer supply line 16supplies oxidizer to the main combustion chamber injector head 22. Afterthe completion of step 228, the method 200 continues to step 232.

At step 232, the preburner combustion chamber exhaust products are mixedand combusted with the oxidizer to produce thrust, the thrust directedthrough the nozzle 24 of the LRE 100. After the completion of step 232,the method 200 ends.

The benefits of the disclosed LRE are achieved by ensuring highreliability, compact design and efficiency of combustion withafterburning of the gas in a combustion chamber at a high level ofpressures, which produces a high specific impulse. Experimentallyconfirmed theoretical calculations show that for combustion with oxygen,methane fuel exceeds the efficiency of kerosene in terms of specificimpulse by around 20 s and is significantly better in cooling capacity.The temperature of combustion products in the combustion chamber islower by 200 K at the optimum oxidizer ratio with methane fuel insteadof kerosene fuel.

The advantages of the indicated fuel—methane (liquefied natural gas(LNG) of various compositions) enable effective and highly reliable LREsof a new concept: Fuel-Rich Staged Combustion. Typically, in a kerosenefuel engine, this is not possible due to heavy soot generation from thefuel-rich combustion. The rationale of this possibility is thatoxygen-methane combustion with an excess of methane will generateminimal or no soot. In this case, the gas temperature before the turbinecan be increased up to approximately 1300 K, which significantlyimproves the performance of fuel-rich oxygen-methane gas. If all otherconditions are equal, this makes it possible to raise the pressure inthe combustion chamber with an increase of the engine's specificimpulse. At the same time, as it is known from the practice ofdevelopment of LREs with afterburning of fuel-rich gas, such gas is notexposed to the danger of ignition in the turbine duct, which is muchmore likely for high-pressure oxidizing gas, increasing system safetyover an oxidizer-rich embodiment. Thus, small damages of the fuel-richgas duct will not lead to an accident or explosion. For a potentialemergency situation, a special protection system can be envisaged in theLRE, which will promptly shutdown the faulty engine without damaging thetest bench, launch complex, or payload.

An important circumstance in the peculiarities of working chemicalprocesses of the disclosed oxygen-methane LRE is that it allows residualpropellant to quickly evaporate from propellant lines after the engine'sshutdown. At the same time, it is well known that kerosene must beremoved from the supply line by force (by a special blowdown purgeprocess), which requires special equipment and reserve of additionalworking fluids on board—this increases a mass of launch vehicle anddecreases a mass of payload. These circumstances significantly reducethe cost of manufacturing an oxygen-methane LRE (estimated 30-40%manufacturing cost reduction versus an oxygen-kerosene). At the sametime, the reliability of the engine increases, since after control andtechnological tests, there is no need to reassemble the hardware. Thus,it is easier and cheaper to reuse the conceptual oxygen-methane ingeneral. Also, the limitations on multiple in-space starts of methaneLRE are removed (which is peculiar to oxygen-kerosene). High reliabilityof oxygen-methane will ensure its long-lasting service life and launchvehicle in-flight safety, which will make it possible to reuse theengine and significantly reduce the cost of payload injection into thenear and far space orbits.

It should be noted that oxygen-methane engines also operate with reducedenvironmental pollution relative to oxygen-kerosene LREs.

An embodiment with a multi-shaft TPU may be realized, however this wouldincrease dry mass of the system.

The start-up of the LRE 100 may be described as follows. Like the methodof FIG. 2 , other methods of use are possible, to include a sequence ofsteps different than those described, a sequence with additional steps,and a sequence with fewer steps.

Start-up of the oxygen-methane LRE is initiated by a starting turbineinstalled at the TPU shaft with cartridge pressure accumulators,numbering equal to the number of the planned system starts.

The propellants are supplied from the propellant tanks (liquid methaneand liquid oxygen) to the inlet devices of the pumps 2 and 6.Afterwards, propellants are supplied through the corresponding lines andcontrol fittings to engine consumers (3-5 and 7-9). Through controlthrottle 10, a part of liquid oxygen that is gasified due to the coolingof gas duct 21 and cooling manifold 13 (cooling jacket of combustionchamber 23), is supplied into the annular fuel-rich preburner chamber18. A small part of liquid oxygen is supplied to the nozzle throat areavia cooling manifold three 15 where it is injected into the lower partmain combustion chamber 22 for film and/or transpiration cooling of thenozzle throat. The residual part of liquid oxygen is supplied to theinjector cavity of main chamber 16, and after spraying it is supplied tothe head of combustion chamber 22 for mixing with the fuel-rich gasgenerated in the annular preburner integrated into the structure ofcombustion chamber 23 as a single structural unit, for completecombustion.

Liquid methane is supplied along the line with corresponding fittings tothe nozzle cooling circuit 11 of the engine with the limited lengthacceptable for reliable cooling, where the fuel is heated up andcompletely gasified in the throat area. Afterwards the fuel (as thegasified methane) is supplied to the second part of this coolingmanifold 11, the jacket of the annular fuel-rich preburner, providingcooling of the external wall of preburner.

Afterwards the mixing of gaseous methane from the cooling manifold 11and a part of gaseous oxygen from the oxidizer cooling manifold 13 inthe preburner combustion chamber with flow alignment vanes 12,generating a fuel-rich gas with the temperature=(890-1300) K, which isassigned depending on the required output power of TPU driving turbine.

At the preset temperature, the generated fuel-rich gas is suppliedthrough the vanes to the blades of the centrifugal turbine 20, andafterwards through the gas duct 21 it is supplied to the combustionchamber 23 for complete combustion of propellant components, in such wayforming the closed propellant supply system (CPSS).

High-temperature combustion products are expanded in the nozzle 24,generating the engine thrust. Control (thrust throttling) of the LRE isperformed through the propellants bypass systems 4, 8 in response to thecommands of the onboard automatic control system (ACS). Shutdown of theLRE is conducted by actuation of shut-off valves and provides multiplerestarts of the engine with a preset number of starts (N) during oneflight of launch vehicle.

In one embodiment, propellants used may include liquid oxygen and RP-1(Rocket Propellant-1) kerosene.

In one embodiment, one or more components of the system re fabricated byway of 3-d printing.

In one embodiment, the LRE may operate with any fuel, to include withoutlimitation methane and kerosene, known to those skilled in the art.

The above embodiments may, in combination or separately, may utilizecomputer software and/or computer hardware (to include, for example,computer-readable mediums) for any of several functions such asautomated control and state estimation, and furthermore may utilize oneor more GUIs for human interaction with modules or elements orcomponents.

Examples of the processors as described herein may include, but are notlimited to, at least one of Qualcomm® Snapdragon® 800 and 801, Qualcomm®Snapdragon® 610 and 615 with 4G LTE Integration and 64-bit computing,Apple® A7 processor with 64-bit architecture, Apple® M7 motioncoprocessors, Samsung® Exynos® series, the Intel® Core™ family ofprocessors, the Intel® Xeon® family of processors, the Intel® Atom™family of processors, the Intel Itanium® family of processors, Intel®Core® i5-4670K and i7-4770K 22 nm Haswell, Intel® Core® i5-3570K 22 nmIvy Bridge, the AMD® FX™ family of processors, AMD® FX-4300, FX-6300,and FX-8350 32 nm Vishera, AMD® Kaveri processors, Texas Instruments®Jacinto C6000™ automotive infotainment processors, Texas Instruments®OMAP™ automotive-grade mobile processors, ARM® Cortex™-M processors,ARM® Cortex-A and ARM926EJ-S™ processors, other industry-equivalentprocessors, and may perform computational functions using any known orfuture-developed standard, instruction set, libraries, and/orarchitecture.

The exemplary systems and methods of this disclosure have been describedin relation to a liquid rocket engine. However, to avoid unnecessarilyobscuring the present disclosure, the preceding description omits anumber of known structures and devices. This omission is not to beconstrued as a limitation of the scopes of the claims. Specific detailsare set forth to provide an understanding of the present disclosure. Itshould however be appreciated that the present disclosure may bepracticed in a variety of ways beyond the specific detail set forthherein.

Furthermore, while the exemplary aspects, embodiments, and/orconfigurations illustrated herein show the various components of thesystem collocated, certain components of the system can be locatedremotely, at distant portions of a distributed network, such as a LANand/or the Internet, or within a dedicated system. Thus, it should beappreciated, that the components of the system can be combined in to oneor more devices or collocated on a particular node of a distributednetwork, such as an analog and/or digital telecommunications network, apacket-switch network, or a circuit-switched network. It will beappreciated from the preceding description, and for reasons ofcomputational efficiency, that the components of the system can bearranged at any location within a distributed network of componentswithout affecting the operation of the system. For example, the variouscomponents can be located in a switch such as a PBX and media server,gateway, in one or more communications devices, at one or more users'premises, or some combination thereof. Similarly, one or more functionalportions of the system could be distributed between a telecommunicationsdevice(s) and an associated computing device.

Furthermore, it should be appreciated that the various links connectingthe elements can be wired or wireless links, or any combination thereof,or any other known or later developed element(s) that is capable ofsupplying and/or communicating data to and from the connected elements.These wired or wireless links can also be secure links and may becapable of communicating encrypted information. Transmission media usedas links, for example, can be any suitable carrier for electricalsignals, including coaxial cables, copper wire and fiber optics, and maytake the form of acoustic or light waves, such as those generated duringradio-wave and infra-red data communications.

Also, while the flowcharts have been discussed and illustrated inrelation to a particular sequence of events, it should be appreciatedthat changes, additions, and omissions to this sequence can occurwithout materially affecting the operation of the disclosed embodiments,configuration, and aspects.

A number of variations and modifications of the disclosure can be used.It would be possible to provide for some features of the disclosurewithout providing others.

In yet another embodiment, the systems and methods of this disclosurecan be implemented in conjunction with a special purpose computer, aprogrammed microprocessor or microcontroller and peripheral integratedcircuit element(s), an ASIC or other integrated circuit, a digitalsignal processor, a hard-wired electronic or logic circuit such asdiscrete element circuit, a programmable logic device or gate array suchas PLD, PLA, FPGA, PAL, special purpose computer, any comparable means,or the like. In general, any device(s) or means capable of implementingthe methodology illustrated herein can be used to implement the variousaspects of this disclosure. Exemplary hardware that can be used for thedisclosed embodiments, configurations and aspects includes computers,handheld devices, telephones (e.g., cellular, Internet enabled, digital,analog, hybrids, and others), and other hardware known in the art. Someof these devices include processors (e.g., a single or multiplemicroprocessors), memory, nonvolatile storage, input devices, and outputdevices. Furthermore, alternative software implementations including,but not limited to, distributed processing or component/objectdistributed processing, parallel processing, or virtual machineprocessing can also be constructed to implement the methods describedherein.

In yet another embodiment, the disclosed methods may be readilyimplemented in conjunction with software using object or object-orientedsoftware development environments that provide portable source code thatcan be used on a variety of computer or workstation platforms.Alternatively, the disclosed system may be implemented partially orfully in hardware using standard logic circuits or VLSI design. Whethersoftware or hardware is used to implement the systems in accordance withthis disclosure is dependent on the speed and/or efficiency requirementsof the system, the particular function, and the particular software orhardware systems or microprocessor or microcomputer systems beingutilized.

In yet another embodiment, the disclosed methods may be partiallyimplemented in software that can be stored on a storage medium, executedon programmed general-purpose computer with the cooperation of acontroller and memory, a special purpose computer, a microprocessor, orthe like. In these instances, the systems and methods of this disclosurecan be implemented as program embedded on personal computer such as anapplet, JAVA® or CGI script, as a resource residing on a server orcomputer workstation, as a routine embedded in a dedicated measurementsystem, system component, or the like. The system can also beimplemented by physically incorporating the system and/or method into asoftware and/or hardware system.

Although the present disclosure describes components and functionsimplemented in the aspects, embodiments, and/or configurations withreference to particular standards and protocols, the aspects,embodiments, and/or configurations are not limited to such standards andprotocols. Other similar standards and protocols not mentioned hereinare in existence and are considered to be included in the presentdisclosure. Moreover, the standards and protocols mentioned herein, andother similar standards and protocols not mentioned herein areperiodically superseded by faster or more effective equivalents havingessentially the same functions. Such replacement standards and protocolshaving the same functions are considered equivalents included in thepresent disclosure.

The present disclosure, in various aspects, embodiments, and/orconfigurations, includes components, methods, processes, systems and/orapparatus substantially as depicted and described herein, includingvarious aspects, embodiments, configurations embodiments,sub-combinations, and/or subsets thereof. Those of skill in the art willunderstand how to make and use the disclosed aspects, embodiments,and/or configurations after understanding the present disclosure. Thepresent disclosure, in various aspects, embodiments, and/orconfigurations, includes providing devices and processes in the absenceof items not depicted and/or described herein or in various aspects,embodiments, and/or configurations hereof, including in the absence ofsuch items as may have been used in previous devices or processes, e.g.,for improving performance, achieving ease and/or reducing cost ofimplementation.

The foregoing discussion has been presented for purposes of illustrationand description. The foregoing is not intended to limit the disclosureto the form or forms disclosed herein. In the foregoing DetailedDescription for example, various features of the disclosure are groupedtogether in one or more aspects, embodiments, and/or configurations forthe purpose of streamlining the disclosure. The features of the aspects,embodiments, and/or configurations of the disclosure may be combined inalternate aspects, embodiments, and/or configurations other than thosediscussed above. This method of disclosure is not to be interpreted asreflecting an intention that the claims require more features than areexpressly recited in each claim. Rather, as the following claimsreflect, inventive aspects lie in less than all features of a singleforegoing disclosed aspect, embodiment, and/or configuration. Thus, thefollowing claims are hereby incorporated into this Detailed Description,with each claim standing on its own as a separate preferred embodimentof the disclosure.

Moreover, though the description has included description of one or moreaspects, embodiments, and/or configurations and certain variations andmodifications, other variations, combinations, and modifications arewithin the scope of the disclosure, e.g., as may be within the skill andknowledge of those in the art, after understanding the presentdisclosure. It is intended to obtain rights which include alternativeaspects, embodiments, and/or configurations to the extent permitted,including alternate, interchangeable and/or equivalent structures,functions, ranges or steps to those claimed, whether or not suchalternate, interchangeable and/or equivalent structures, functions,ranges or steps are disclosed herein, and without intending to publiclydedicate any patentable subject matter.

What is claimed is:
 1. A liquid rocket engine (LRE) comprising: a fuelsupply line containing a fuel; an oxidizer supply line containing anoxidizer; an oxidizer throttle receiving the oxidizer from the oxidizersupply line, the oxidizer supply line splitting into a first partialoxidizer supply line, a second partial oxidizer supply line, and a thirdpartial oxidizer supply line; a turbine mounted to a turbine shaft alonga longitudinal centerline of the LRE at a longitudinal proximal locationof the LRE, the turbine rotating with the turbine shaft about thelongitudinal centerline; a gas duct in fluid communication with theturbine; a main combustion chamber having a main chamber injector headin fluid communication with both the gas duct and the second partialoxidizer supply line; a nozzle positioned longitudinally distal to themain combustion chamber and in fluid communication with the maincombustion chamber; a preburner combustion chamber positioned axiallydistal to the main combustion chamber and having a preburner injectorhead in fluid communication with both a first cooling manifold and asecond cooling manifold, wherein the preburner combustion chamber formsan annulus around the main combustion chamber; the first coolingmanifold in fluid communication with the fuel supply line; the secondcooling manifold positioned axially between the preburner combustionchamber and the main combustion chamber, the second cooling manifold influid communication with the first partial oxidizer supply line; and athird cooling manifold in fluid communication with the third partialoxidizer supply line and a nozzle throat of the nozzle; wherein: thefuel received by the first cooling manifold from the fuel supply linereceives thermal energy from the nozzle and is supplied to the preburnerinjector head; the oxidizer received by the second cooling manifold fromthe first partial oxidizer supply line receives thermal energy from atleast one of the gas duct, the main combustion chamber, and thepreburner combustion chamber, the oxidizer supplied to the preburnerinjector head; the preburner injector head injects the fuel with theoxidizer to create preburner combustion exhaust products which flow fromthe preburner combustion chamber to the turbine and drive the turbineabout the turbine shaft, the preburner combustion exhaust productsflowing from the turbine to the gas duct; and the main combustionchamber injector head injects the preburner combustion exhaust productsreceived from the gas duct with oxidizer received from the secondpartial oxidizer supply line to produce an LRE thrust directed throughthe nozzle exit.
 2. The liquid rocket engine of claim 1, wherein thefuel is methane.
 3. The liquid rocket engine of claim 1, wherein themain combustion chamber comprises a main chamber ignitor which ignitesthe preburner combustion exhaust products with the oxidizer.
 4. Theliquid rocket engine of claim 1, wherein the turbine is a centrifugalturbine.
 5. The liquid rocket engine of claim 1, further comprising aset of turbine guide vanes operating to direct the preburner combustionexhaust products into the turbine.
 6. The liquid rocket engine of claim1, wherein the gas duct is positioned axially between the turbine andthe main combustion chamber.
 7. The liquid rocket engine of claim 1,wherein at least some of the oxidizer becomes a gasified oxidizer due tothermal energy transfer from at least one of the gas duct, the maincombustion chamber, and the preburner combustion chamber.
 8. The liquidrocket engine of claim 1, wherein at least some of the fuel becomes agasified fuel due to thermal energy transfer from the nozzle to the atleast some of the fuel, the gasified fuel supplied to the preburnerinjector head.
 9. A liquid rocket engine (LRE) comprising: a fuel supplyline containing a fuel; an oxidizer supply line containing an oxidizer;an oxidizer throttle receiving the oxidizer from the oxidizer supplyline, the oxidizer supply line splitting into a plurality of partialoxidizer supply lines; a turbine mounted to a turbine shaft along alongitudinal centerline of the LRE at a longitudinal proximal locationof the LRE, the turbine rotating with the turbine shaft about thelongitudinal centerline; a main combustion chamber having an injectorhead in fluid communication with the turbine and at least one of theplurality of partial oxidizer supply lines; a nozzle positionedlongitudinally distal to the main combustion chamber and in fluidcommunication with the main combustion chamber; and a preburnercombustion chamber positioned axially distal to the main combustionchamber and having a preburner injector head, the preburner combustionchamber forming an annulus around the main combustion chamber; wherein:the preburner injector head injects a gasified fuel with a gasifiedoxidizer to create preburner combustion exhaust products which flow fromthe preburner combustion chamber to the turbine and drive the turbineabout the turbine shaft, the preburner combustion exhaust productsflowing from the turbine to the main combustion chamber; and the maincombustion chamber injector head injects the preburner combustionexhaust products received from the turbine with oxidizer to produce anLRE thrust directed through the nozzle exit.
 10. The liquid rocketengine of claim 9, further comprising a gas duct positioned axiallybetween the turbine and the main combustion chamber.
 11. The liquidrocket engine of claim 10, wherein at least some of the oxidizer becomesa gasified oxidizer due to thermal energy transfer from at least one ofthe gas duct, the main combustion chamber, and the preburner combustionchamber.
 12. The liquid rocket engine of claim 11, wherein at least someof the fuel becomes a gasified fuel due to thermal energy transfer fromthe nozzle to the at least some of the fuel, the gasified fuel suppliedto the preburner injector head.
 13. The liquid rocket engine of claim 9,wherein the fuel is methane.
 14. The liquid rocket engine of claim 9,wherein the turbine is a centrifugal turbine.
 15. The liquid rocketengine of claim 9, further comprising a set of turbine guide vanesoperating to direct the preburner combustion exhaust products into theturbine.
 16. A method of operating a liquid rocket engine (LRE)comprising: providing a LRE comprising: a fuel supply line containing afuel; an oxidizer supply line containing an oxidizer; an oxidizerthrottle; a turbine mounted to a turbine shaft along a longitudinalcenterline of the LRE at a longitudinal proximal location of the LRE; amain combustion chamber having an injector head in fluid communicationwith the turbine and at least one of the plurality of partial oxidizersupply lines; a nozzle positioned longitudinally distal to the maincombustion chamber and in fluid communication with the main combustionchamber; and a preburner combustion chamber positioned axially distal tothe main combustion chamber and having a preburner injector head, thepreburner combustion chamber forming an annulus around the maincombustion chamber; receiving the oxidizer from the oxidizer supplyline; splitting the oxidizer supply line into a plurality of partialoxidizer supply lines; injecting, using the preburner injector head, agasified fuel with a gasified oxidizer to create preburner combustionexhaust products which flow from the preburner combustion chamber to theturbine; rotating the turbine about the turbine shaft; flowing thepreburner combustion exhaust products from the turbine to the maincombustion chamber; and injecting, using the main combustion chamberinjector head, the preburner combustion exhaust products received fromthe turbine with oxidizer to produce an LRE thrust directed through thenozzle exit.
 17. The method of claim 16, the LRE further comprising agas duct positioned axially between the turbine and the main combustionchamber, wherein at least some of the oxidizer becomes a gasifiedoxidizer due to thermal energy transfer from at least one of the gasduct, the main combustion chamber, and the preburner combustion chamber.18. The method of claim 16, wherein at least some of the fuel becomes agasified fuel due to thermal energy transfer from the nozzle to the atleast some of the fuel, the gasified fuel supplied to the preburnerinjector head.
 19. The method of claim 16, wherein: the turbine is acentrifugal turbine; and the fuel is methane.
 20. The method of claim16, the LRE further comprising a set of turbine guide vanes operating todirect the preburner combustion exhaust products into the turbine.